High performance gas turbine engines typically rely on increasing turbine inlet temperatures to increase both fuel economy and overall power ratings. These higher temperatures, if not compensated for, oxidize engine components and decrease component life. Component life has been increased by a number of techniques. Said techniques include internal cooling with air bled from an engine compressor section. Bleeding air results in efficiency loss however. In addition, stationary gas turbine engines typically may have less available compressed air than moving gas turbine engines.
U.S. Pat. No. 3,806,274 issued to Moore on Apr. 23, 1974 shows a gas turbine blade with a hollow interior space which is divided to form flow passages for cooling medium. In particular, the flow passages are bounded by the sides of a sheet-like insert between the two blade walls. Fins extend between the insert and the blade walls. The fins commence at one end of the blade and extend in a spiral-like path around the opposite sides of the insert. The insert is located between a large number of pimples and by a series of helical fins cast onto the interior surfaces of the blade walls. The insert stops short of both the leading and trailing edges of the blade, thus leaving spaces around which air may pass in order to progress from one side of the insert to the other.
The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors.